Tiedown ply for reducing core crush in composite honeycomb sandwich structure

ABSTRACT

We reduce core crush in honeycomb sandwich structure by using a peripheral tiedown ply, generally in combination with a scrim-reinforced barrier film, between the composite laminate and the core along the panel chamfer to prevent slipping of the barrier film and outer laminates relative to the core during curing. We produce superior panels with lighter weights, improved mechanical properties, and more predictable structural performance.

NOTICE OF GOVERNMENT RIGHTS

The present invention was made during performance of Contract No.F33657-91-C-0006 awarded by the Air Force. The Government has certainrights in the invention.

TECHNICAL FIELD

The present invention relates to composite honeycomb sandwich structurehaving improved resistance to core crush. In a preferred embodiment, weadhere resin impregnated fabric sheets forming outer, opposed, skinsurfaces to a chamfered honeycomb core, optionally, with an intermediatebarrier film to eliminate resin flow from the skins to the core. Weinterleave a peripheral tiedown ply between the core and skins along thechamfer to reduce core crush.

BACKGROUND ART

Aerospace honeycomb core sandwich panels (having composite laminateskins cocured with adhesives to the core through autoclave processing)find widespread use today because of the high stiffness-to-weight (i.e.,"specific stiffness) and strength-to-weight (i.e., specific strength)ratios the panels afford. Typical honeycomb core sandwich panels aredescribed in U.S. Pat. No. 5,284,702; 4,622,091; and 4,353,947, which weincorporate by reference. Alteneder et al., Processing andCharacterization Studies of Honeycomb Composite Structures, 38th Int'lSAMPE Symposium, May 10-13, 1993 (PCL Internal No. 200-01/93-AWA)discusses common problems with these panels, including core collapse(i.e., core crush), skin laminate porosity, and poor tool surfacefinish. We incorporate this article by reference.

As Hartz et al. described in U.S. Pat. No. 5,604,010 entitled "CompositeHoneycomb Sandwich Structure," with a high flow resin system, largeamounts of resin can flow into the core during the autoclave processingcycle. Such flow robs resin from the laminate, introduces a weightpenalty in the panel to achieve the desired performance, and forces overdesign of the laminate plies to account for the flow losses. The resinloss from the laminate plies also reduces the thickness of the curedplies which compromises the mechanical performance. To achieve thedesired performance and the corresponding laminate thickness, additionalplies are necessary with resulting cost and weight penalties. Becausethe weight penalty is severe in terms of the impact on vehicleperformance and cost in modern aircraft and because the flow is arelatively unpredictable and uncontrolled process, aerospace design andmanufacture dictates that flow into the core be eliminated orsignificantly reduced. In addition to the weight penalty from resin flowto the core, we discovered that microcracking that originated in themigrated resin could propagate to the bond line and degrade mechanicalperformance. Such microcracking potential poses a catastrophic threat tothe integrity of the panel and dictates that flow be eliminated or, atleast, controlled.

Flow from the laminates to the core occurs because of viscosityreduction of the resin (i.e., thinning) at the elevated processingtemperatures. Therefore, prior art attempts to solve the flow problemhave generally focused on retaining the ambient temperature viscosity ofthe resin at the curing temperatures. For example, one might alter theprocessing cycle to initiate curing of the resin during a slow heat-up,low pressure step to induce resin chain growth before high temperature,high pressure completion. In this staged cure cycle, one would try toretain the resin's viscosity by building molecular weight at lowtemperatures. Higher molecular weight resins have inherently higherviscosity so they remain thicker and are resistant to damaging flow tothe core. Unfortunately, with a staged cure cycle, too much flow stilloccurs, and the potential problems of microcracking still abound. Also,facesheet porosity might increase beyond acceptable limits. Furthermore,a modified cure cycle increases autoclave processing time. Increasedprocessing time translates to a significant fabrication cost increasewith risk of rejection of high value parts at the mercy of uncontrolledand inadequately understood factors. We incorporate the Hartz et al.U.S. Pat. No. 5, 604,010 by reference.

U.S. Pat. No. 5,445,861 describes composite sandwich structure for soundabsorption (acoustic insulation) and other applications. The sandwichstructures have seven layers as follows:

(1) an outer skin;

(2) a small celled honeycomb or foam core;

(3) a frontside inner septum;

(4) a large celled middle honeycomb core;

(5) a backside, inner septum;

(6) a backside, small celled honeycomb or foam core; and

(7) an inner skin.

Tuned cavity absorbers in the middle honeycomb core absorb sound.Performance of this structure suffers from resin flow to the cells ofthe honeycomb cores during fabrication for the reasons already discussedand because such flow alters the resonance of the structure. Weincorporate this patent by reference.

The Hartz et al. process of U.S. Pat. No. 5,604,010 eliminates resin(matrix) flow into the honeycomb core for sandwich structure using highflow resin systems and results in reproducibility and predictability insandwich panel fabrication and confidence in the structural performanceof the resulting panel. Hartz et al. use a scrim-supported barrier filmbetween the fiber-reinforced resin composite laminates and the honeycombcore. This sandwich structure is lighter for the same performancecharacteristics than prior art panels because the resin remains in thelaminate (skin) where it provides structural strength rather thanflowing to the core where it is worthless, introducing excess weight andpotential panel failure. Hartz et al. also generally use an unsupportedfilm adhesive between the barrier film and the laminates to bond thelaminates to the barrier film. With these layers (which might becombined into one product), they achieved improved performance, retainedthe resin in the laminates and thereby reduced excess resin thatdesigners otherwise needed to design into the panels to account forresin flow into the core, and reliably fabricated panels in which theyhad structural confidence.

We discovered that core crush frequently occurred in the chamfer regionof honeycomb core when we cured a panel having a scrim-supported barrierfilm, particularly when we tried to use lighter weight core materials.We subsequently discovered that we could reduce core crush in thesepanels by including a tiedown ply in contact with the core beneath thebarrier film (and adhesive) because the tiedown ply reduced slippage ofthe barrier film relative to the core during curing.

SUMMARY OF THE INVENTION

Our invention relates to a method for reducing core crush in compositehoneycomb sandwich structure, especially panels of the general typeHartz et al. described. We incorporate one or more tiedown plys incontact with the core in its chamfer regions around the periphery toeliminate slippage of the skin over the core during autoclave curing,and, thereby, to eliminate core crush that results from the movement.

Our invention also relates to the resulting composite honeycomb sandwichstructure. There, we usually minimize the weight by trimming theinternal area of the tiedown ply(s) so that it frames and slightlyoverlaps the chamfer of the underlying core. By controlling coreslippage, we are able to use the lighter density honeycomb core toproduce structures without costly scrap due to core crush. We reducemanufacturing costs both by saving time, materials, and rework/scrap andby improving the reliability of the manufacturing process to produceaerospace-quality panels having the highest specific strength andspecific stiffness.

The tiedown ply also provides a path for egress of volatiles from thecore and to equalize the pressure which permits us to maintain thecorrect pressures within the core to further reduce core crush.

BRIEF DESCRIPTION OF THE DRAWINGS.

FIG. 1 illustrates a typical composite honeycomb sandwich structure.

FIG. 2 is a schematic, partial sectional view of the skin-core interfacein Hartz-type sandwich structure having a scrim-supported barrier filmto prevent resin flow from the skin to the core.

FIG. 3 is a schematic, partial sectional view of prior art honeycombsandwich structure, suffering resin flow to the core, using a supportedfilm adhesive without a barrier film.

FIG. 4 is another schematic, partial sectional view showing sandwichstructure with resin depletion in the skin, but where the resin isprevented from reaching the core with a bulging, unsupported barrierfilm.

FIG. 5 is a schematic, sectional elevation showing core crush of ahoneycomb sandwich panel caused by core and barrier film slippage.

FIG. 6 is another schematic, sectional elevation showing the use of atiedown ply to reduce core crush.

DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT

Before discussing our method to eliminate core crush, we will initiallydescribe typical composite honeycomb sandwich structure.

A Hartz-type composite honeycomb sandwich panel of U.S. patentapplication Ser. No. 08/587,160 now U.S. Pat. No. 5,604,010 minimizes,eliminates, or significantly reduces resin flow from the laminates tothe core, thereby permitting a simpler processing cycle that is morerobust for the manufacture of aerospace structure. Such a sandwich panel100 (FIG. 1) generally has outer facesheets or skins 102 adhered to acentral honeycomb core 106. The finished skins 102 comprise laminates oflayers of fiber-reinforced organic matrix resin in a cured andconsolidated composite form. The core 106 can be paper, synthetic paper,metal, composite, or the like, as appropriate for the application. Inpanels of the present invention, we obtain higher specific strengths andhigher specific stiffnesses because we reduce core crush duringautoclave curing by incorporating at least one tiedown ply between thecore 106 and skin 102 to reduce damaging slippage between the core andskin that otherwise often occurs.

To prevent flow of resin from the composite laminate skin to the core,Hartz et al. use an unsupported film adhesive 108 (FIG. 2), a barrierfilm 110, and a scrim-supported film adhesive 112 between the skin 102and the core 106 to keep resin out of the cells 114 of the core 106.

FIG. 3 illustrates the core-filling problems that can result when a filmadhesive 112 is used alone without the barrier film 110 and filmadhesive 108. Cells 114 of the honeycomb fill with resin 118 whichmigrates from the laminates and which thereby depletes the resin in theskin 102. Resin depletion impacts structural performance because itreduces ply thickness. Resin depletion increases total weight since thecell resin 118 is simply waste. In all cases, uncontrolled resin flowand depletion makes the panel suspect, especially to microcracking thatcan begin in the cell resin 118 during thermal cycling and migrate tothe fiber-reinforced skin 102, especially at the bond line between theskin 102 and core 106.

FIG. 4 illustrates undesirable bulging that can occur if a barrier film110 is used without a scrim-supported film adhesive 112 to try toeliminate cell resin 118. Here, a waste resin bulge 120 protrudesdownwardly into the cells 114 of the honeycomb core 106. While the resinis contained in the bulge 120, the skin 102 is still depleted in resin.The flow of resin to bulge 120 imposes structural performance and weightpenalties comparable to the uncontrolled condition illustrated in FIG.3.

As shown in FIG. 2 with the film adhesive 108, barrier film 110, andscrim-supported film adhesive 12, resin flow is checked without cellresin 118 or resin bulges 120. We discovered, however, that the barrierfilm produced a slip plane between the laminate skins and the core whichoften resulted in core crush during the autoclave processing cycle. In22 of 31 test panels, in fact, Hartz et al. experienced core crush intheir initial trials. This rate of failure is unacceptable from a costand schedule perspective. Our tiedown plys in the chamfer region reducethe frequency of or eliminate damaging core slippage and the core crushattributable to such slippage.

For bismaleimide laminated skins made with RIGIDITE® 5250-4-W-IM7GP-CSW,RIGIDITE® 5250-4-W-IM7-GP-CSX, and RIGIDITE® 5250-4-WIM7-GP-PW prepregfrom Cytec Engineered Materials, Inc. (Cytec), the film adhesive 108preferably is 0.015 psf METLBOND® 2550U adhesive, also available fromCytec. The film adhesive provides additional resin to promote a qualitybond between the laminate and barrier film 110. The barrier film 110preferably is a 0.001 inch thick, bondable grade, surface treatedKAPTON® polyimide barrier film capable of withstanding the cure cycle toprovide a resin impermeable membrane between the skin 102 and core 106.The scrim preferably is fiberglass, "Style 104" fiber cloth and the filmadhesive 112 is 0.06 psf METLBOND® 2550G adhesive, available from Cytec.The scrim-supported film adhesive prevents the barrier film from bulginginto the core cells, thereby retaining the resin in the laminate (i.e.,skin layers) so that the cured ply thickness is maximized and thereby,we achieve maximum performance at minimum weight for the panels.

The film adhesive 108, barrier film 110, and film adhesive 112 can bepurchased as a single item from Cytec as METLBOND® 2550B-0.082 36 ".

The plies of the skin 102 typically are prepregs of carbon fiberimpregnated with bismaleimide thermoset resin, although the presentinvention applies to other resin systems. Tows might be used in place ofthe prepreg. The film adhesive 108 should be tailored to achieve anadequate bond between the skin 102 and barrier film 110. The honeycombcore generally is HRP Fiberglass Reinforced Phenolic honeycomb availablefrom Hexcel.

The supported film adhesive and barrier film layers in the sandwichstructure also function as corrosion barriers between the skin 102 andcore 106 in the case where the core is metal, such as aluminum, and theskin includes a galvanically dissimilar material, such as carbon fiber.

Additional information concerning preferred panels is presented in thetechnical paper: Hartz et al., "Development of a BismaleimadelCarbonHoneycomb Sandwich Structure," SAMPE, March, 1996, which we incorporateby reference. This paper describes both the barrier film improvement andthe tiedown ply method of the present invention.

The Hartz-type panels provide mechanical and physical edgebandproperties equivalent to solid BMI/carbon laminate (cured at 0.59 MPa(85 psig)). Our tests confirm that in our panels the edgebandcured-ply-thickness is equivalent to a solid laminate and that theedgeband 160 (FIGS. 5 & 6) met the requirements of the solid laminatenondestructive inspection specification. The edgeband and facesheetmechanical performance improved over results we achieved with sandwichstructure lacking the scrim-supported adhesive, barrier film, adhesivecombination. The flatwise tensile mechanical performance also met designrequirements.

Preconditioning the core to eliminate volatile evolution during curingby heating the core to about 235° C. (455° F.), prior to laying up thesandwich panel, especially for phenolic core, eliminates core-laminatedisbonding otherwise caused by outgassing from the core. The tiedownplies, in addition, provide egress channels for escape of volatiles fromthe core and for pressure equalization.

Having described a preferred, Hartz-type composite honeycomb sandwichpanel, we will now turn to describing the improved panels of the presentinvention and their method of manufacture.

Core crush 200 (FIG. 5) occurs in the chamfer region 155 when thebarrier film 110 and core 106 slip relative to the facesheets 102 whenautoclave pressure is applied and when the resin is melted. As shown inFIG. 5, the barrier films 100 and core 106 have moved toward the rightto compress the core in the chamfer region 155 to produce the core crush200. The skin 102 has sagged in the edgeband region 160 where the coremoved away.

Referring now to FIG. 6, our improved honeycomb sandwich panel includesat least one tiedown ply 150 in contact with the core 106 along achamfer 155. Such a chamfer (i.e. an angled transition in the core,often at the edgeband 160) typically occurs around the periphery of thepanel, but it might also occur intermediate of the panel at join linesor hardpoints where fasteners or passthroughs might be necessary in theassembled structure.

Typically we use a single ply 150 of carbon fiber or fiberglass fabricwith a conventional 0/90, degree fiber orientation in the fabrication ofbismaleimide panels having 5 or 8 Ib/ft³ HRP core, like Hartz et al.describe. The tiedown ply 150 functions to prohibit or to limit slippageof the skin relative to the core so as to reduce core crush otherwiseattributable to the slippage. The tiedown ply 150 anchors the core withthe inherent roughness of the fabric when the preform is heated duringthe autoclave processing cycle and the matrix resin softens, melts, and,for high flow resins, essentially liquefies. With these panels, we cansave between 2.5-4 lb/ft³ of core because we can use lighter densityhoneycomb core without suffering core crush. For a fighter, this changecan save as much as 25 lbs per vehicle.

As shown in FIG. 6, the tiedown ply 150 is a narrow, peripheral stripthat contacts the core 106 along at least a portion of the chamfer 155for about 1 inch overlap with the core 106 and extends outward into theedgeband 160 beyond the trimline 165 of the part. The tiedown ply 150might be on either the flat side of the chamfer or the angled surface(which is how we show it in FIG. 6). The key factor is that the tiedownply 150 contact the core beneath the adhesive and barrier film 110 whichis used to bond the laminate skin to the core. The tiedown ply 150 iscutaway everywhere in the body of the part other than a narrowperipheral area in the chamfer region, and forms a peripheral framearound the edge of the panel. In this way, the tiedown ply 150 allows anadhesive interface between the core 106 and the skins 102 in the panelregion.

Traditionally, we use four complete cover sheet tiedown plies 175 in aneffort to anchor the layers and the core, and we show all these plies inFIG. 6. These traditional plies 175 were commonly used in sandwich panelfabrication prior to introducing the Hartz-type barrier film, and wecommonly use them all, although we believe we can now eliminate all butthe outer plies and the peripheral, core contacting tiedown ply 150.That is, we would use three total plies rather than five, as FIG. 6shows.

The tiedown plies 150 and 175 extend through the edgeband 160 beyond thenet trim line 165 to anchoring points that we tape to the layup mandrel.To further prevent slippage of the tiedown plies, we have incorporated alow curing (i.e. 121° C. for BMI panels) film adhesive 180 between thetiedown plies just outside the net trim line of the part. The filmadhesive 180 eliminates movement of one ply relative to the others whenwe apply pressure during the autoclave curing cycle.

Thus, the tiedown method of the present invention can save material,reduce cost, and save weight, if only the "picture frame" peripheraltiedown ply 150 is used (with the traditional, internal sheets omitted).The normal tiedown procedure entails plys on the outer surfaces of theskins and internally between the skin and underlying adhesive. Thistiedown system fails without the "picture frame" ply because the barrierfilm 110 permits the core to slip.

For lightweight core (i.e. 5-8 lb/ft³) with the bismaleimide prepreg andadhesive system previously described, we hold the chamfer angle to20°±2°.

By "chamfer" we mean an angled, cut region of the honeycomb coretapering from full thickness to no thickness with a steady slope. Achamfer is used at the edge band of a composite honeycomb sandwich panelto provide a smooth transition between the structural body of the panelthat has the embedded honeycomb and a connecting edge band lacking anyhoneycomb core. The method of the present invention allows us to usemuch steeper chamfer angles than traditional practices often require ifone is to avoid core crush without one tiedown ply. While we prefer a20° chamfer, we believe that we could increase the angle to whateverangle suited the panel design requirements.

By "autoclave processing" we mean the cycle of elevated temperature andpressure applied to the panel to consolidate and cure resin in thelaminate while bonding or otherwise adhering the cured laminate to thehoneycomb core.

If core crush occurs, the damage to the panel is generally so extensivethat repair is impossible so the part is scraped. The cost of today'sadvanced composite resins and reinforcing fibers requires a process thatvirtually eliminates core crush. Otherwise, the processing costs areprohibitive. With panels being designed as close to the design edge aspossible, core crush is a significant issue. The method of the presentinvention solves cores crush concerns at the root cause.

While we have described preferred embodiments, those skilled in the artwill readily recognize alterations, variations, and modifications, whichmight be made without departing from the inventive concept. Therefore,interpret the claims liberally with the support of the full range ofequivalents known to those of ordinary skill based upon thisdescription. The examples are given to illustrate the invention and arenot intended to limit it. Accordingly, define the invention by theclaims and limit the claims only as necessary in view of the pertinentprior art.

We claim:
 1. Composite honeycomb sandwich structure having improvedresistance to core crush, comprising:(a) a honeycomb core, havingopposed faces, core cells, and a peripheral chamfer along the margin ofthe core; (b) at least one composite laminate on each face, the laminatehaving plies of fiber-reinforced matrix resin adhered to the core toform a sandwich structure; (c) a barrier film between the laminate andthe core to bond the laminate and core and to eliminate resin flow fromthe laminate into the core cells; and (d) a tiedown ply, having a wovenfabric impregnated with resin, in contact with the chamfer of the corebeneath the barrier film, and extending to a trim margin of the part forbeing tied down to eliminate slippage of the barrier film relative tothe core and, in so doing, to reduce core crush.
 2. The structure ofclaim 1 wherein the laminate includes bismaleimide matrix resin.
 3. Thestructure of claim 2 wherein the adhesive includes bismaleimide.
 4. Thestructure of claim 3 further comprising a film adhesive layer betweenthe barrier film and each laminate.
 5. The structure of claim 1 whereinthe barrier film is a polyimide.
 6. The structure of claim 1 furthercomprising a supporting scrim between the barrier film and the core toprevent sagging of the barrier film into the core cells.
 7. Thestructure of claim 1 wherein the tiedown ply includes a carbon fiber orfiberglass fabric.
 8. The structure of claim 1 wherein the tiedown plyis cutaway in all but the region of the core defined by the chamfer tominimize weight.
 9. The structure of claim 1, wherein the tiedown plyhas a 0/90 degree fiber orientation.
 10. Composite honeycomb sandwichstructure resistant to core crush caused by slippage of a compositelaminate along a chamfer of a honeycomb core, comprising:(a) a honeycombcore having a chamfer; (b) a tiedown ply of resin-impregnated wovenfabric contacting the chamfer; and (c) at least one laminate contactingthe tiedown ply at the chamfer and adhered to the core through thetiedown plywherein the tiedown ply prevents damaging slippage of thelaminate relative to the core that would produce core crush duringautoclave curing of the structure.
 11. The structure of claim 10 whereinthe tiedown ply includes a carbon fiber or fiberglass fabric. 12.Composite honeycomb sandwich structure having improved resistance tocore crush, comprising:(a) a honeycomb core, having core cells and aperipheral chamfer; (b) at least one composite laminate having plies offiber-reinforced bismaleimide matrix resin adhered to the core; (c) abondable grade polyimide barrier film between the laminate and the coreto eliminate resin flow from the laminate into the core cells; (d) anadhesive between the barrier film and the core to bond the laminate andcore; (e) a supporting scrim between the adhesive and the core toprevent sagging of the barrier film into the core cells; and (f) atiedown ply of bismaleimide resin-impregnated woven fabric in contactwith the chamfer of the core beneath the adhesive and scrim to eliminateslippage of the barrier film relative to the core and, in so doing, toreduce cor crush.
 13. The structure of claim 12, wherein the tiedown plyhas a 0/90 degree fiber orientation.
 14. Composite honeycomb sandwichstructure resistant to core crush caused by slippage of a compositelaminate along a chamfer of a honeycomb core, comprising:(a) a honeycombcore having a chamfer; (b) a tiedown ply of resin-impregnated wovenfabric contacting the chamfer; and (c) at least one laminate contactingthe tiedown ply at least at the chamfer and adhered to the corewhereinthe tiedown ply includes a carbon fiber or fiberglass fabric andprevents damaging slippage of the laminate relative to the core thatwould produce core crush during autoclave curing of the structure. 15.The structure of claim 14, wherein the tiedown ply has a 0/90 degreefiber orientation and is cutaway in all regions of the core other thanthe chamfer.